Field of the Invention
The present invention relates to a turbine blade that includes at least one cooling passage formed in the blade so as to extend in a blade height direction and that has blade surfaces covered with thermal barrier coating.
Background Art
A gas turbine is a kind of rotary machine and has been used for a power source mainly for propulsion of an aircraft or power generation. A gas turbine is provided with a compressor, a combustor, and a turbine. The compressor draws in and compresses air and generates compressed air. The combustor combusts fuel with the compressed air generated by the compressor and generates high-temperature combustion gas. The turbine is rotated by the combustion gas (mainstream gas) generated by the combustor.
For an improvement in performances of a gas turbine, there is a requirement for increasing the temperature of the combustion gas. However, a problem (specifically, creeping, thinning due to oxidation, or the like) in which an increase in temperature of the combustion gas causes damage to a turbine blade (that is, a stator blade or a rotor blade) easily occurs. As a countermeasure for this problem, there is a method of forming a cooling passage inside the turbine blade and causing cooling air to flow through the cooling passage. There is also a method of covering a blade surface (that is, a surface of a blade material) with thermal barrier coating.
Although an increase in the thickness of the thermal barrier coating brings about a stronger effect of thermally shielding the blade surface from the high-temperature mainstream gas, an aerodynamic performance of the turbine blade deteriorates. Thus, the thickness of the thermal barrier coating at a blade trailing edge gradually reduces toward aback side according to JP-A-2013-194667. In doing so, the blade trailing edge width reduces and thus improves the aerodynamic performance.
Detailed description will be given of JP-A-2013-194667. According to JP-A-2013-194667, a design point on a suction side is set at a position of a tailing end of a final cooling passage, which is the closest to a blade trailing edge in at least one cooling passage extending in a blade height direction (specifically, at a position through which a straight line that passes through the tailing end of the final cooling passage and is perpendicular to a camber line passes through the blade surface on the suction side), on the blade surface on the suction side in each blade section perpendicular to the blade height direction. In addition, a thickness distribution of the thermal barrier coating on the suction side of each blade section is configured such that the thickness of the thermal barrier coating is uniform from a blade leading edge to the design point on the suction side and gradually reduces from the design point on the suction side toward the backside up to the blade trailing edge.
Similarly, a design point on a pressure side is set at a position of the tailing end of the final cooling passage (specifically, at a position through which the straight line that passes through the tailing end of the final cooling passage and is perpendicular to the camber line passes through a blade surface on the pressure side) on the blade surface on the pressure side in each blade section perpendicular to the blade height direction. In addition, the thickness distribution of the thermal barrier coating on the pressure side of each blade section is configured such that the thickness of the thermal barrier coating is uniform from the blade leading edge to the design point on the pressure side and gradually reduces from the design point on the pressure side toward the back side up to the blade trailing edge.